Gas turbine engines are generally known in the art and used in a wide range of applications, such as aircraft engines and auxiliary power units for aircraft. In a typical configuration, turbine engines include rows of stator vanes and rotor blades disposed in an alternating sequence along the axial length of a generally annular hot gas flow path. The rotor blades are mounted on rotor platforms at the periphery of one or more platforms of rotor disks that are coupled in turn to a main engine shaft. Hot combustion gases are delivered from an engine combustor to the annular hot gas flow path, thus resulting in rotary driving of the rotor disks to provide an engine output.
The stator vanes and rotor blades typically have arcuate airfoil shapes with generally concave pressure sides and generally convex suction sides extending axially in chords between opposite leading and trailing edges. During operation, the aerodynamic contours of the stator vanes and rotor blades, and corresponding flow passages therebetween, are configured in an attempt to maximize energy extraction from the combustion gases. The complex three-dimensional (3D) configuration of the stator vanes and rotor blades varies radially in span along the airfoils as well as axially along the chords of the airfoils between the leading and trailing edges. As such, the velocity and pressure distributions of the combustion gases over the various surfaces, as well as within the corresponding flow passages, can vary.
Undesirable pressure losses in the combustion gas flow paths correspond with undesirable reduction in overall turbine efficiency. One common source of turbine pressure losses is the formation of horseshoe vortices generated as the combustion gases are split in their travel around the leading edges of the rotor blades. Particularly, a pair of counterrotating passage vortices are formed at the leading edge of the rotor blades. These vortices travel along the opposite pressure and suction sides of each rotor blade and behave differently due to the different pressure and velocity distributions therealong. For example, computational analysis indicates and flow testing supports that the pressure side vortex migrates away from the endwall toward the trailing edge and then interacts with the suction side vortex flowing aft thereto. The interaction of the pressure and suction side vortices occurs near the midspan region of the airfoils and may create total pressure loss and a corresponding reduction in turbine efficiency. As the pressure side vortex breaks away from the rotor blade, a corner vortex also develops. These corner vortices, as well as the passage vortices, can disrupt film cooling air along the platform surface, and reduce the cooling effectiveness thereof. More film cooling holes may therefore be required to improve cooling performance, which in turn increases cooling air requirements and decreases turbine efficiency.
Indeed, from the viewpoint of efficiency, it is desirable to operate the turbine at temperatures as high as possible. As a practical matter, however, the complexity of the vortices complicates the mechanisms for cooling the components, particularly the rotor platform. Thus, in order to economically produce turbines capable of sustained high temperature operation, other schemes to increase cooling effectiveness and efficiency are necessary.
Accordingly, it is desirable to provide an improved gas turbine engine assembly that suppresses vortex formation that may otherwise lead to reduced efficiency. In addition, it is desirable to provide an improved gas turbine engine assembly that replenishes any cooling film that has been disrupted as a result of these vortices. Furthermore, other desirable features and characteristics of the present invention will become apparent from the subsequent detailed description of the invention and the appended claims, taken in conjunction with the accompanying drawings and this background of the invention.